The present disclosure generally relates to alumina protective coatings for thermal barrier coatings, and more specifically, thermally sprayed alumina-based protective coatings for thermal barrier coatings utilized in various surfaces on gas turbine components.
In turbine engines, such as but not limited to aircraft and power generation turbines, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The air-fuel mixture is combusted, and the resulting hot combustion gas is passed through a turbine mounted on the same shaft. The turbine includes a rotor with turbine blades supported on its periphery, and a stationary (that is, not rotating) gas turbine flow path shroud that confines the combustion gas to flow through the annulus between the rotor and the shroud, and thence against the turbine blades. The constrained flow of hot combustion gas turns the turbine by contacting an airfoil portion of the turbine blade, which turns the rotor shaft and provides power to the compressor. The rotating turbine blades and the gas turbine stationary flow path shroud are heated to high temperatures by the hot combustion gas.
To prevent these components from getting too hot, thermal barrier coatings (TBCs) are often coated onto various surfaces of the turbine components to help insulate the components from the high temperatures in the hot gas path. TBCs are an increasingly important element in current and future gas turbine engine designs because of the higher operating temperatures in gas turbine engines. Examples of turbine engine parts and components for which such thermal barrier coatings are desirable include turbine blades and vanes, turbine shrouds, buckets, nozzles, combustion liners and deflectors, and the like. These thermal barrier coatings are deposited onto a metal substrate (or more typically onto a bond coat layer on the metal substrate for better adherence) from which the part or component is formed to reduce heat flow and to limit the operating temperature these parts and components are subjected to. This metal substrate typically comprises a metal alloy such as a nickel-, cobalt-, and/or iron-based alloy (e.g., a high temperature super alloy).
The thermal barrier coating is usually prepared from a ceramic material, such as a chemically (metal oxide) stabilized zirconia. Examples of such chemically stabilized zirconias include yttria-stabilized zirconia, scandia-stabilized zirconia, calcia-stabilized zirconia, magnesia-stabilized zirconia, and combinations thereof. The thermal barrier coating of choice is typically an yttria-stabilized zirconia ceramic coating. A representative yttria-stabilized zirconia thermal barrier coating usually comprises about 7-weight % yttria and about 93-weight % zirconia. The thickness of the thermal barrier coating depends upon the metal part or component it is deposited on, but is usually in the range of from about 3 to about 70 mils thick for high temperature gas turbine engine parts.
Although significant advances have been made in improving the durability of thermal barrier coatings for turbine engine components, such coatings are still susceptible to various types of damage, including objects ingested by the engine, erosion, oxidation, and attack from environmental contaminants. In addition, in trying to achieve reduced thermal conductivity, other properties of the thermal barrier coating can be adversely impacted. For example, the composition and crystalline microstructure of a thermal barrier coating, such as those prepared from yttria-stabilized zirconia, can be modified to impart to the coating an improved reduction in thermal conductivity, especially as the coating ages over time. However, such modifications can also unintentionally interfere with desired spallation resistance, especially at the higher temperatures that most turbine components are subjected to. As a result, the thermal barrier coating can become more susceptible to damage due to the impact of, for example, objects and debris ingested by the engine and passing through the turbine sections thereof. Such impact damage can eventually cause spallation and loss of the thermal barrier coating.
In addition, at the higher temperatures of engine operation, the environmental contaminants can adhere to the heated or hot thermal barrier coating and subsequently cause damage. For example, environmental contaminants can form compositions that are liquid or molten at the higher temperatures at which gas turbine engines operate. These molten contaminant compositions can dissolve the thermal barrier coating, or can infiltrate its porous structure, i.e., can infiltrate the pores, channels, or other cavities in the coating. Upon cooling, the infiltrated contaminants solidify and reduce the coating strain tolerance, thus initiating and propagating cracks that cause delamination, spalling and loss of the thermal barrier coating material either in whole or in part. Damage may also result from the freezing contaminants having a different coefficient of thermal expansion relative to the TBC.
These pores, channels or other cavities that are infiltrated by such molten environmental contaminants can be created by environmental damage, or even the normal wear and tear that results during the operation of the engine. However, the porous structure of pores, channels or other cavities in the thermal barrier coating surface more typically is the result of the processes by which the thermal barrier coating is deposited onto the underlying bond coat layer-metal substrate. For example, thermal barrier coatings that are deposited by air plasma spray techniques tend to create a sponge-like porous structure of open pores in at least the surface of the coating. By contrast, thermal barrier coatings that are deposited by physical (e.g., chemical) vapor deposition techniques tend to create a porous structure comprising a series of columnar grooves, crevices or channels in at least the surface of the coating. This porous structure can be important in the ability of these thermal barrier coatings to tolerate strains occurring during thermal cycling and to reduce stresses due to the differences between the coefficient of thermal expansion (CTE) of the coating and the CTE of the underlying bond coat layer/substrate.
For turbine engine parts and components having thermal barrier coatings with such porous surface structures; environmental contaminant compositions of particular concern are those containing oxides of calcium, magnesium, aluminum, silicon, and mixtures thereof. These oxides combine to form contaminant compositions comprising mixed calcium-magnesium-aluminum-silicon-oxides (Ca—Mg—Al—SiO), hereafter referred to as “CMAS.” During normal engine operations, CMAS can become deposited on the TBC surface, and can become liquid or molten at the higher temperatures of normal engine operation. Damage to the TBC typically occurs when the molten CMAS infiltrates the porous surface structure of the thermal barrier coating. After infiltration and upon cooling, the molten CMAS solidifies within the porous structure. This solidified CMAS causes stresses to build within the thermal barrier coating, leading to partial or complete delamination and spalling of the coating material and, thus partial or complete loss of the thermal protection provided to the underlying metal substrate of the part or component.
CMAS mitigation coatings are often needed for gas turbine operation above 2200° F. Many turbine engines are operating in this temperature regime. Without CMAS mitigation, the TBC is often compromised and the component could fail before it next service interval.
Alumina is known to protect TBCs from CMAS infiltration. It has also demonstrated ability to yield smooth anti-fouling surfaces for hot gas path components. Current processes for depositing alumina include chemical vapor deposition (CVD) for CMAS protection and the application of alcohol slurry solutions for anti-fouling applications. However, CVD processes are prohibitively expensive and slow. The CVD process does not lend its self to the large scale of gas turbine components since it requires controlled atmospheric conditions. With regard to the use of alcohol based slurry solutions, these solutions generally provide a limited shelf life, high cost and are typically incorporate hazardous organic solvents. The slurry coating process also requires a post coating drying and furnace curing. The slurry coatings are not practical for CMAS mitigation because the thicknesses of the coatings are limited. Moreover, with regard to anti-fouling applications, the cohesive strength of the slurry coating is limited because it cannot be sintered without melting the metallic component to which it is applied. Another limitation of the current slurry process is that it is limited to service temperatures of 2000° F. making it incompatible with the CMAS environment. Moreover, current slurry methods also generally require a 12 hour dry followed by 8 hour high temperature cure, making manufacturing costs high.
Accordingly, it would be desirable to provide improved methods for protecting the thermal barrier coating to the adverse effects of such environmental contaminants. In particular, a need exists to protect such thermal barrier coatings that can be suitably used to mitigate both CMAS and fouling of hot path gas surfaces.